The invention is related to the field of space launch vehicles for use in launching a payload from a stationary ground-based position into orbit, and specifically to a cryogenic propellant depletion system for maximizing the utilization of cryogenic propellant by the booster stage of the space launch vehicle.
Launch vehicles are generally used to launch payloads, such as satellites or scientific equipment, from the Earth""s surface into space. Launch vehicles generally include one or more rocket engines arranged to fire at different times, or stages, as the launch vehicle travels from the earth""s surface into orbit. The different stages are fired sequentially and typically include at least a first stage or xe2x80x9cboosterxe2x80x9d stage and a second upper stage. The booster stage is designed to launch and deliver the payload a predetermined distance above the earth before exhaustion. Upon exhaustion, the booster stage and upper stage separate whereupon the upper stage is fired the to deliver the payload the remainder of the distance into a desired orbit. In the case where the booster stage is a reusable component, the booster stage controllably falls back to the earth""s surface upon separation for retrieval, refitting, and future use.
The booster stage""s rocket engine(s) typically utilize liquid propellants and in the case of bi-propellant rockets generally include two or more propellant tanks, booster pumps, a combustion chamber, plumbing interconnecting the various components, and a nozzle for accelerating and/or discharging the combustion product. Liquid propellant rockets generally utilize a liquid fuel such as RP-1 (i.e., kerosene) and an oxidizer such as liquid oxygen (LOX), which are stored in separate propellant tanks and brought into contact in the combustion chamber to provide thrust.
In order to preserve the booster stage rocket engine for re-use, the engine must be properly shut down at or near the end of the launch boost stage. That is, the engine must be shut down prior to violating any engine requirements that may result in some sort of permanent engine or vehicle damage. For, example, reusable rocket engines generally require that the propellant(s) be supplied with a minimum xe2x80x9cheadxe2x80x9d pressure in order for the engine to properly function. As the propellant in one or both of the propellant tanks nears exhaustion, the head pressure generally drops. This head pressure drop may potentially result in engine and/or booster pump damage if the head pressure drops below a minimum allowable threshold (i.e., engine requirement). Further, proper engine shut down generally requires a predetermined mass of one or both of the propellants to prevent engine and/or booster pump damage which my result in catastrophic failure. Therefore, it is important to initiate engine shut down prior to the propellant(s) dropping below any minimum allowable threshold, in both reusable and expendable booster rockets. However, initiating engine shut down to prevent engine damage and/or rocket failure must be balanced with the desire to fully utilize the available propellant to maximize the launch vehicle""s booster stage performance. As will be appreciated, it is desirable to utilize the propellant right up until the last possible moment prior to exceeding a minimum allowable threshold in order to maximize boost. Accordingly, it is desirable to continuously monitor the amount of the propellant remaining such that engine shut down may be initiated just prior to the propellant dropping below any minimum allowable threshold.
Existing propellant monitoring systems generally utilize hot point sensors which indicate a transition between liquid to gas in the propellant supply system through, for example, a change in a monitored capacitance. In these systems booster engine cut off (BECO) is initiated a predetermined time after the propellant level drops below one of the hot point sensors. Unfortunately, it is often difficult to precisely determine the amount of the remaining propellant using hot point sensors when a two-phase mixture of the propellant and/or an ullage gas exist in the system. This is especially true with cryogenic propellants which are susceptible to xe2x80x9cboiling offxe2x80x9d (i.e., liquid oxygen to gaseous oxygen) and which may introduce a two-phase mixture into the cryogenic storage tank and the plumbing interconnecting the storage tank to the rest of the system. Two-phase mixtures of the cryogenic propellants make determination of the remaining propellant parameters difficult as hot point sensors may prematurely or belatedly indicate the remaining level/pressure of propellant or otherwise provide erratic signals. As a result of inaccurate propellant parameter measurements the booster stage engine may prematurely shut down and fail to utilize all available propellant, thus, reducing booster stage performance. Alternatively, the propellant may be depleted beyond a minimum allowable threshold required for proper engine shutdown prior to initiating engine shut down which may result in violation of the engine requirements and/or engine damage.
In view of the foregoing, a primary object of the present invention is to provide a system for monitoring a cryogenic propellant, irrespective of that propellant being a pure liquid. Another objective of the present invention is to provide a cryogenic depletion monitoring system that is easily adaptable for operation with current launch vehicles.
One or more of the above-noted objectives, as well as additional advantages, are provided by the present invention, which includes a cryogenic depletion monitoring system for use in monitoring a cryogenic propellant in a feed line between a cryogenic storage tank and a booster engine in a space launch vehicle. The cryogenic depletion monitoring system utilizes a processing system, memory, supported logic, and one or more pressure sensor readings to generate one or more parameters related to the propellant in the feed line. The propellant parameters may then be used to determine when to initiate booster engine cut off such that booster stage performance is maximized while no engine requirements are violated.
According to a first aspect of the present invention, a cryogenic propellant depletion monitoring system is provided that includes a processing system containing logic to determine and monitor at least a first propellant parameter associated with the cryogenic propellant in the launch vehicle""s cryogenic feed line. Additionally, the system contains control means to initiate engine shut off once at least one of the determined propellant parameters falls below a minimum allowable threshold. The system is operable to determine these propellant parameters and initiate booster engine shutdown irrespective of the cryogenic propellant being a pure liquid.
Various refinements exist of the features noted in relation to the subject first aspect of the present invention. Further features may also be incorporated into the subject first aspect of the present invention as well. These refinements and additional features may exist individually or in any combination. For example, the processing system and the supported logic may utilize any available information for determining the propellant parameters associated with the propellant in the feed line. This information may include, inter alia, algorithms, material constants and/or prior test data that may be stored in an accessible memory structure.
In a preferred embodiment of the present invention, the processing system utilizes at least two pressure measurements associated with the cryogenic system to determine the one or more parameters related to the propellant in the feed line. These pressure measurements may include a first pressure measurement of the propellant from a point along the length of the feed line and a second pressure measurement from the cryogenic storage tank. In this regard, the processing system may utilize the pressure measurements to determine a pressure differential between these two points in the cryogenic system. Preferably, a first set of one or more measurements is taken near the bottom of the feed line and a second set of one or more pressure measurements is taken in the cryogenic propellant storage tank such that a pressure differential across may be calculated across at least a portion of the feed line.
The feed line pressure measurement is indicative of the pressure head at the bottom of the feed line. However, in the case where pressurized ullage gas is utilized to expel the cryogenic propellant from the storage tank, the pressure head will include an ullage gas pressure component. In this regard, the propellant logic is configured to subtract this ullage pressure from the feed line pressure head to produce an xe2x80x9ceffectivexe2x80x9d feed line pressure head that represents the head pressure supplied by the propellant independent of ullage pressure. As will be appreciated, this effective feed line pressure head includes the propellant that is in a pure liquid state and propellant that is in a two phase mixture state within the feed line, thus providing a system wherein a feed line pressure head is calculable irrespective of the cryogenic propellant being a pure liquid.
The processing system and its supported logic may further utilize this effective feed line pressure differential (i.e., effective head pressure) to calculate the usable mass of the propellant within the feed line. In this regard, the processing system may utilize one or more additional parameters to determine the feed line mass. For example, to calculate the feed line propellant mass, the physical constraints of the feed line such as length, area and/or volume must be known as well as the density of the cryogenic propellant, which may vary with temperature necessitating propellant temperature to be known. Additionally, in order to calculate a usable mass of the liquid oxygen propellant in the feed line, the effective feed line pressure head may be normalized for launch vehicle acceleration. That is, the effective feed line pressure head may be divided by the current acceleration to eliminate the effects of acceleration on the pressure head and, thus, provide a xe2x80x9cnormalizedxe2x80x9d pressure head. This normalized pressure head may then be used to determine the mass of the cryogenic propellant in the feed line.
In order to produce a usable feed line propellant mass, any unusable mass contained a two-phase region within the feed line must be taken into account. In this regard, the processing system may operatively access stored information for determining the usable percentage of feedline propellant. For example, experimental test results from earlier flights or ground tests may be used to determine usable mass percentages compared to a total mass in the feed line. The processor is operative to use this information, such as a predetermined correction factor or xe2x80x9ccurvexe2x80x9d, to determine the usable cryogenic propellant mass contained in the feed line.
The processing system is further operative to monitor the determined feed line propellant parameters. In this regard, the processing system is operative to determine minimum allowable requirements for each monitored feed line propellant parameter and then continuously compare the feed line parameters to these minimum allowable requirements. For example, a minimum head pressure requirement may be required at the bottom of the feed line for proper engine operation. In this regard, the processing system may calculate a current minimum required feed line pressure and then compare the effective feed line pressure head to this value. Further, the processing system may simultaneously compare another propellant parameter to another corresponding minimum allowable requirement. Once any of the monitored requirements falls below a minimum allowable threshold, the processing system is operative to issue a command to the controller, which in turn initiates shutdown of the booster engine.
According to a second aspect of the present invention, a method for monitoring cryogenic propellant in a feed line interconnecting a propellant storage tank to a booster engine in a space launch vehicle is provided. The method includes the steps of taking a first pressure measurement associated with the feed line, taking a second pressure measurement associated with the propellant storage tank and utilizing the first and second pressure measurements to determine a propellant pressure differential in the feed line. Finally, the method includes utilizing the pressure differential to calculate at least a first additional propellant parameter associated with the propellant in the feed line.
Various refinements exist of the features noted in relation to the subject second aspect of the present invention. Further features may also be incorporated into the subject second aspect of the present invention as well. These refinements and additional features may exist individually or in any combination. For instance, the method of the second aspect could also include taking multiple pressure readings at one or more points associated with the feed line and/or storage tank to provide increased measurement accuracy and a fail safe redundancy system.
Preferably, the pressure measurements associated with the feed line are taken near the bottom of the feed line while the pressure measurement associated with the propellant storage tank is taken from the propellant storage tank""s ullage space. This latter measurement allows determination of the ullage gas pressure, which may then be subtracted from the feed line pressure measurement to determine an effective feed line pressure differential independent of ullage pressure. As will be appreciated this effective feed line pressure differential represents the head pressure supplied by pure liquid propellant and two phase propellant in the feed line.
The step of utilizing the pressure differential to calculate at least a first additional propellant parameter, in a preferred embodiment, includes utilizing the effective pressure head to calculate a usable propellant mass in the feed line. In order to calculate the usable feed line mass, the effective pressure head is normalized by dividing by acceleration to provide a pressure head independent of acceleration. This normalized pressure head is then converted to a propellant mass utilizing one or more stored constants to account for unusable propellant mass contained in a two-phase that may be expected under a given set of conditions. These conditions may include, inter alia, mass flow rate, pressure head and/or geometry of the feed line. In this regard, correction factors determined from prior flights and/or experiments may be utilized to calculate a usable percentage of the remaining propellant mass in the feed line.
The method of the second aspect of the present invention may further comprise comparing the pressure differential and/or an additional propellant parameter, such as usable propellant mass, to corresponding minimum allowable thresholds related to engine requirements. Accordingly, the method includes the step of initiating booster engine shutdown when at least one of the pressure differential and/or the first additional propellant parameter falls below its corresponding minimum allowable threshold. As will be appreciated, this provides a method for monitoring a cryogenic propellant in a feed line that allows the propellant to be utilized up to the last possible moment prior to violating a minimum allowable engine requirement, which may result in permanent engine damage and/or failure.
According to a third aspect of the present intention, a cryogenic propellant depletion monitoring system is provided for monitoring at least a first parameter related to a cryogenic propellant in a feed line interconnecting a propellant storage tank and a booster engine in a space launch vehicle. The system comprises at least a first pressure determining device for determining at least a first pressure associated with the feed line and a processing platform configured to receive and utilize this first pressure for determining at least one parameter associated with the cryogenic propellant in the feed line. In order to determine the cryogenic propellant parameter(s), the processing platform is further operable to accessing a memory structure that contains information and/or instructions for use in producing the propellant parameters. For example, the memory structure may contain pertinent material constants related to the cryogenic propellant as well as protocols/algorithms for use in calculating one or more propellant parameters. Finally, the system contains a controller in communication with processing platform for initiating booster engine shutdown when one of the determined propellant parameters drops below a predetermined value.
The system""s pressure determining device(s) may be any device operable to provide the desired pressure measurement associated with the feed line, such as a pressure transducer. In a preferred embodiment a plurality of pressure measurement devices are utilized to provide a feed line pressure. In order to provide consistency amongst the plurality of pressure measurement devices, these devices must measure the feed line pressure at the same point. In this regard, the preferred embodiment utilizes a manifold in fluid communication with the feed line. That is, the manifold is interconnected to the feed line while the plurality of pressure determining devices are interconnected to the manifold. As will be appreciated, the pressure within the manifold is equal to the pressure at the feed line connection point. In this regard, the plurality of pressure determining devices interconnected to the manifold each effectively provide a measurement at the same point along the feed line. The plurality of pressure measurements allows for increased measurement accuracy as well as a fail safe redundancy system. For example, the plurality of measurements may be averaged or center selected to provide a more accurate pressure measurement.
The system may utilize a second pressure determining device to determine a pressure associated with the storage tank, which may also be utilized in determining propellant parameters associated with the cryogenic propellant. Again, any suitable pressure sensor may be utilized to provide the storage tank pressure. However, the launch vehicle may utilize a constant pressurization system for the storage tank where the pressure within the storage tank remains a predetermined constant value. Accordingly, in this embodiment, a separate pressure determining device will not be utilized, rather, the known storage tank pressure may be stored within the memory structure for use by the processing platform as needed.
The system""s processing platform could be a single processing device or a group of inter-operational processing devices so long as the processing system is operable to perform the desired functions for producing the propellant parameters. In this regard, the processing platform may be configured within an existing processor of the space launch vehicle. That is, an existing computer, such as a flight control computer, may provide the necessary functionality for the processing platform. Alternatively, the processing platform may be a stand alone unit configured to independently produce the propellant parameters. Irrespective of the platform used, the processing platform is operatively connected to a memory structure containing the necessary instructions which can be retrieved and executed by the processing platform to produce the desired propellant parameters. The memory structure may be any appropriate storage media such as disks, tapes, integrated circuits, etc.
The pressure measurement devices may be directly interconnected to the processing platform through an interface associated with the processing platform. That is, the pressure measurement devices may be electrically interconnected to, for example, an analog to digital converter which receives analog signals from the devices and provides digital outputs of these signals to the processing platform for use in determining the propellant parameters.